1.Constructed the wing of the selected aircraft using the different type of materials. 2.Created Solid works simulation numerical model and perform analysis. 3.Conducted the strength calculations according to strength, stiffness and stability Requirements. 4.Prepared work Drawings.

1. 1
PROJECT REPORT
Naga Manikanta Kommanaboina
Master Science – Aeronautical Engineering
Kaunas University of Technology

2. 2
CONTENTS
1.CONSTRUCT THE WING OF THE SELECTEDAIRCRAFT
USING COMPOSITE METRIALS
1.1Describe the purpose and characteristics of the……………………. 3
desired aircraft wing.
1.2Composite a calculation scheme with all possible…………………. 4
load cases and in accordance with the requirements
of the certification Specification (cs…..…)
2. CREAT SOLID WORKS SIMULATION NUMERICAL
MODEL AND PERFORM ANALYSIS.
2.1 Build solid works 3D model………………………………………. 9
2.2 Conduct strength calculations according to strength,
stiffness and stability Requirements………………………………….. 10
2.3 Prepare work drawing……………………………………………...16
CONCLUSION……………………………………………………….18
REFERENCES……………………………………………………….19

3. 3
1. CONSTRUCT THE WING OF THE SELECTEDAIRCRAFT
USING COMPOSITE METRIALS
1.1 Describe the purpose and characteristics of the desired aircraft
wing
For my fighter jet I choose the tapered wing
Tapered Wing
These types of wing have a different chord at root and different chord at tip.
This wing will
perform well in supersonic cases. They are the swept wings; there are two kinds
of swept wings namely,
Swept Forward
Swept Backward
Swept forward wing is highly unstable wing and structurally weak.
The swept backward wing is suits for Supersonic cases.
I need the more stable and structurally strong that’s why i selected the swept
back wing.
This wing will operate well in both subsonic and supersonic.
The swept back wings are delaying the shock waves and accompanying
aerodynamic drag rise caused by fluid compressibility near the speed of sound,
improving performance.
The characteristics of the desired aircraft wing is
Wing area = 53.38 m^2

4. 4
Wing span = 12.26 m
Root chord = 6.803m
Tip chord = 1.90m
Wing aerodynamic chord = 4.35
Aspect ratio = 2.81
Sweep angle = 73.52 º
1.2 Composite a calculation scheme with all possible load cases and in
accordance with the requirements of the certification specification
(cs…)
Takeoff weight =18159.54Kg
climb weight = 18159.54kg
WING DESCRIPTION
Lift varies along the wing span due to the variation in chord length, angle of
attack and Sweep along the
span. Schreck’s curve defines this lift distribution over the wing span of an
Aircraft, also called simply
as Lift Distribution Curve.
Schrenk’s Curve is given by
Y=Y1+Y2 / 2
Where
y1 is Linear Variation of lift along semi wing span also named as L1

5. 5
y2 is Elliptic Lift Distribution along the wing span also named as L2
LINEAR LIFT DISTRIBUTION
By Schrenk’s Curve,
Lift at Root,
L_root = ρ×V2×Cl×Croot =0.13668×(614.46)2×1.45×6.803 / 2
Lroot = 254487.8935N/m
Similarly, Ltip = 71075.55N/m
LOAD ESTIMATION ON WING
Self-Weight,
Mwing
Mto = 0.25
Wwing = 18159.54 × 9.81 = 44536.27N
Wport wing = 22268.135N
WStarboard wing = -22268.135N
FUEL WEIGHT
Wfuel per wing = 33562.46Kg
For 2 Wings= 67124.92 Kg
Using general formula for straight,
Line y=mx+c
dy = 1.5325 m
Dy= 5250.42 – 1.5325 m
m = 1427.1 N/m2
yf = 1427.10x-5250.42

6. 6
Area of Ellipse,
X=4a/3π =4(21.60)/3π
X= 9.167
Trapezoidal,
b/3 = 2.7585/3 = 0.9195 =0.9195 + 0.4195
= 1.839 m
Fig1: Fuel load distribution

7. 7
Fig2: Total weight distribution
Flight load
(i)Sec. 23.321 - General.
(a) Flight load factors represent the ratio of the aerodynamic force component
(acting normal to the assumed longitudinal axis of the airplane) to the weight
of the airplane. A positive flight load factor is one in which the aerodynamic
force acts upward, with respect to the airplane.
(b) Compliance with the flight load requirements of this subpart must be
shown—
(1) At each critical altitude within the range in which the airplane may be
expected to operate;
(2) At each weight from the design minimum weight to the design maximum
weight; and
(3) For each required altitude and weight, for any practicable distribution of
disposable load within the operating limitations specified in Secs. 23.1583
through 23.1589.
(c) When significant, the effects of compressibility must be taken into account]
Amdt. 23-45, Eff. 09/07/93
Sec. 23.333 - Flight envelope.
Maneuvering envelope. Except where limited by maximum (static) lift
coefficients, the airplane is assumed to be subjected to symmetrical maneuvers
resulting in the following limit load factors:
(1) The positive maneuvering load factor specified in Sec. 23.337 at speeds up
to VD
(2) The negative maneuvering load factor specified in Sec. 23.337 at VC; and
(3) Factors varying linearly with speed from the specified value at VC to 0.0
at VD for the normal and -1.0 at VD for the acrobatic and utility categories.

8. 8
Sec. 23.337 - Limit maneuvering load factors.
The positive limit maneuvering load factor n may not be less than 6.0 for
acrobatic category airplanes.
Sec. 23.391 - Control surface loads.
The control surface loads specified in Secs. 23.397 through 23.459 are
assumed to occur in the conditions described in Secs. 23.331 through 23.351
HORIZONTAL STABILIZING AND BALANCING SURFACES
Sec. 23.421 - Balancing loads.
A horizontal [surface] balancing load is a load necessary to maintain
equilibrium in any specified flight condition with no pitching acceleration.
Horizontal [balancing] surfaces must be designed for the balancing loads
occurring at any point on the limit maneuvering envelope and in the flap
conditions specified in Sec. 23.345
Sec. 23.425 - Gust loads.
[(a) Each horizontal surface, other than a main wing, must be designed for
loads resulting from-
(1) Gust velocities specified in Sec. 23.333(c) with flaps retracted; and
(2) Positive and negative gusts of 25 f.p.s. nominal intensity at VF ,
corresponding to the flight conditions specified in Sec. 23.345(a)(2). [(b)
[Reserved.]
Sec. 23.423 - Maneuvering loads.
Each horizontal surface and its supporting structure, and the main wing of a
canard or tandem wing configuration, if that surface has pitch control, must be
designed for the maneuvering loads imposed by the a sudden movement of the
pitching control, at the speed VA, to the maximum aft movement, and the

9. 9
maximum forward movement, as limited by the control stops, or pilot effort,
whichever is critical.
VERTICAL SURFACES
Sec. 23.441 - Maneuvering loads.
(a) At speeds up to VA, the vertical surfaces must be designed to withstand
the following conditions. In computing the loads, the yawing velocity may be
assumed to be zero:
(1) With the airplane in unaccelerated flight at zero yaw, it is assumed that the
rudder control is suddenly displaced to the maximum deflection, as limited by
the control stops or by limit pilot forces.
(2) With the rudder deflected as specified in paragraph of this section, it is
assumed that the airplane yaws to the overswing sideslip angle. In lieu of a
rational analysis, an overswing angle equal to 1.5 times the static sideslip angle
of paragraph of this section may be assumed.
2. CREAT SOLID WORKS SIMULATION NUMERICAL
MODEL AND PERFORM ANALYSIS.
2.1 Build soild works 3D model

10. 10
Fig1: 3D solid works wing model
2.2 Conduct strength calculations according to strength, stiffness and
stability Requirements.
ALUMINUM 7075
Fig 2: Stress analysis

11. 11
Fig3: Mesh of the wing
Fig4: Displacement of the wing

12. 12
Fig5: Strain analysis
Fig6: Buckling analysis

13. 13
ALLOY STEEL
Fig7: Static mode of Stress analysis
Fig8: Static displacement of the wing

14. 14
Fig9: Static strain analysis
Fig 10: Buckling analysis

15. 15
STRUCTURAL ANALYSIS OF AIRCRAFT WING
A static analysis calculates the effects of steady loading conditions on a
structure, while ignoring inertia and damping effects, such as those caused by
time varying loads. A static analysis can, however. Include steady inertia loads
(such as gravity and rotational velocity), and time varying loads that can be
approximated as static equivalent loads.
Static analysis determines the displacements, stresses, strains and forces in
structures and components caused by loads that do not induce significant
inertia and damping effect. Steady loading and response conditions are
assumed; that is, the loads and structures response are assumed to vary slowly
with respect to time.
BUCKLING ANALYSIS
Buckling analysis is a technique used to determine buckling loads-critical
loads at which a structure becomes unstable-and buckled mode shapes-the
characteristic shape associated with a structure's buckled response.

16. 16
2.3 PREPARE WORK DRAWINGS
Fig11: Front view of the wing

17. 17
Fig12: Right view of the wing
Fig13: Side view of the wing
Fig14: Distance between the ribs

18. 18
CONCLUSION
In this project, the aircraft wing model was created by the solid works software.
We observed, aircraft wing with Aluminum 7075 obtained load factor -7.09
and deformation is 76.64 and for Alloy steel obtained load factor is -2.32 and
deformation scale is 77. By using of the Aluminum 7075. we can get the more
strength wing and good corrosion resistance is less important. It is known to
have a high strength to weight ratio and certain tempers offers decent resistance
to stress-corrosion cracking. Dynamic pressure on leading edge is decreasing
with increasing the angle of attack. It is observed from the analysis; static
pressure on lower surface is increasing with increasing the angle of attack. We
also conclude that static pressure is increased with increasing the angle of
attack is increased. Dynamic pressure on lower surface is decreasing with
increasing angle of attack whereas; static pressure is increasing on lower
surface. We conclude that dynamic pressure is increased with increasing the
angle of attack is increased.
Constrai
n
Numb
er of
crew
and
passen
ger
Engin
es
Max.
takeoff
weight
(N)
Turn load
/ speed /
height
Rate-
of
climb
at SL
Take off
run /
speed
Cruise
speed /
height
Magnitu
ed
1 2 18159.
54 KG
+9.0-
3.0g
254m
/s
286.88k
m/h
1963k
m/h
Constrai
n
Servic
e
ceiling
AR of
wing
CD
min,
Rolling
drag
coeff
Take
-off
CLT
O
Take-
off
CDTO
Landin
g speed
Magnitu
ed
65,000
ft
2.81 0.284 -
5,658kg/
m
0.8 0.46 95.32m
/s

19. 19
REFERENCES
1.Aircraft performance and design, “John D. Anderson, Jr. University of
Maryland”
2. Aircraft design – A conceptual approach, “Daniel P. Raymer president
Conceptual Research
Cooperation, Sylmar California”
3. An example of airplane preliminary design procedure – Jet Transport, “E.
G. Tulapurkara, A. Venkattraman, V. Ganesh”
4. Aircraft Design A Systems Engineering Approach, “Mohammad H. Satrapy,
Daniel Webster College,
New Hampshire, USA